WO2021199802A1 - Pale statique et turbine à gaz d'aéronef - Google Patents

Pale statique et turbine à gaz d'aéronef Download PDF

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Publication number
WO2021199802A1
WO2021199802A1 PCT/JP2021/007122 JP2021007122W WO2021199802A1 WO 2021199802 A1 WO2021199802 A1 WO 2021199802A1 JP 2021007122 W JP2021007122 W JP 2021007122W WO 2021199802 A1 WO2021199802 A1 WO 2021199802A1
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WO
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Prior art keywords
blade
stationary
fan
wing
stationary blade
Prior art date
Application number
PCT/JP2021/007122
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English (en)
Japanese (ja)
Inventor
真也 楠田
Original Assignee
株式会社Ihi
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by 株式会社Ihi filed Critical 株式会社Ihi
Priority to JP2022511663A priority Critical patent/JP7294528B2/ja
Priority to EP21781375.7A priority patent/EP4130436A4/fr
Publication of WO2021199802A1 publication Critical patent/WO2021199802A1/fr
Priority to US17/808,114 priority patent/US11867090B2/en

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers
    • F04D29/544—Blade shapes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/323—Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/129—Cascades, i.e. assemblies of similar profiles acting in parallel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/301—Cross-sectional characteristics
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/96—Preventing, counteracting or reducing vibration or noise
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This disclosure relates to a stationary blade of an aircraft gas turbine engine and an aircraft gas turbine engine.
  • Patent Document 1 discloses an outlet guide wing intended to reduce the above-mentioned moving and stationary wing interference noise.
  • the moving blade interference sound is generated by the periodic interference between the wake of the moving blade (speed loss region called wake) and the stationary blade (for example, the exit guide blade) provided behind the moving blade. Further, the moving and stationary blade interference noise is generated not only in the fan but also in other rotating machines such as a compressor and a turbine. In general, it is known that the sound pressure level of the moving and stationary wing interference sound increases in proportion to the 6th power of the velocity when the sound source is a dual pole sound source and the 8th power when the sound source is a quadrupole sound source. ing. On the other hand, a decrease in exhaust speed is directly linked to a decrease in thrust. Therefore, it is required to reduce noise while avoiding fluctuations in exhaust speed.
  • the wing body may have the airfoil cross section satisfying the above conditions from the tip side of the wing body to the hub side of the wing body.
  • the code ratio of the maximum blade thickness position on the chip side of the blade body may be larger than the code ratio of the maximum blade thickness position on the hub side of the blade body.
  • the stationary blade according to the present disclosure may be a fan stationary blade.
  • the aircraft gas turbine engine according to the present disclosure accommodates a fan blade, a fan case accommodating the fan blade, and a core engine provided behind the fan blade, and is between the fan case.
  • a core case defining a bypass flow path for the working fluid and a stationary blade according to the present disclosure provided as a fan blade behind the fan rotor blade in the bypass flow path are provided.
  • FIG. 1 is a schematic cross-sectional view of an aircraft gas turbine engine according to an embodiment of the present disclosure.
  • FIG. 2 is a diagram showing a stationary wing according to the embodiment of the present disclosure.
  • FIG. 3 is a diagram showing the relationship between the inclination of the moving blade and the maximum blade thickness position of the stationary blade in the axial direction, and
  • FIG. 3A is a cross-sectional view taken along the line IIIA-IIIA (cross-sectional view on the chip side) shown in FIG.
  • FIG. 3B is a cross-sectional view taken along the line IIIB-IIIB (cross-sectional view on the hub side) shown in FIG.
  • FIG. 4 is a graph showing the distribution of the wing surface Mach number on each dorsal side and each ventral side of the stationary blade according to the embodiment and the conventional stationary blade
  • FIGS. 4 (a), 4 (b) and 4 (C) shows the blade surface Mach number distribution at 90% span, 50% span, and 10% span, respectively.
  • FIG. 5 is a graph showing the numerical analysis results of each sound pressure level of the stationary blade according to the embodiment and the stationary blade of the conventional example.
  • turbofan engine will be given as an example of an aircraft gas turbine engine according to the present embodiment. Further, the turbofan engine is simply referred to as an "engine".
  • the turbofan engine according to the present embodiment may be a geared turbofan engine or another gas turbine engine provided with a fan. In either case, the bypass ratio does not matter.
  • the stationary blade according to the present embodiment is not limited to application to a fan which is a rotary machine (axial flow machine), and other rotary machines (axial flow machine) such as a low pressure compressor, a high pressure compressor, a high pressure turbine, and a low pressure turbine. ) Is also applicable.
  • FIG. 1 is a schematic cross-sectional view of the engine 1 according to the present embodiment.
  • the engine 1 includes a core engine 10 and a fan 20 provided in front of the core engine 10.
  • the core engine 10 includes a low-pressure compressor 11L, a high-pressure compressor 11H, a combustor 12, a high-pressure turbine 13H, a low-pressure turbine 13L, and a core nozzle 14. These are housed in the core case 15 and arranged along the axis 2. In other words, they are arranged from the upstream side (front, left side in FIG. 1) to the downstream side (rear, right side in FIG. 1) of the mainstream of the working fluid (that is, air or combustion gas).
  • the core engine 10 according to this embodiment is a multi-stage turbine engine.
  • the number of stages of the compressor and the turbine may be, for example, the above-mentioned two stages or three stages.
  • the stretching direction of the shaft 2 is defined as the axial direction AD.
  • the circumferential direction centered on the axis 2 is defined as the circumferential direction CD. It is assumed that the rotation direction RD of each moving blade (including the fan moving blade) described later coincides with the circumferential direction CD.
  • the high pressure compressor 11H is provided behind the low pressure compressor 11L.
  • the high-pressure compressor 11H includes a moving blade fixed to the high-pressure shaft 16b and a stationary blade fixed to the outer wall of the high-pressure compressor 11H. Similar to the low-pressure compressor 11L, the stationary blades and the moving blades of the high-pressure compressor 11H are alternately installed along the shaft 2, and both are arranged in the circumferential direction CD.
  • the high-pressure compressor 11H further compresses the working fluid compressed by the low-pressure compressor 11L and supplies it to the combustor 12.
  • a fuel supply system (not shown) is connected to the combustor 12.
  • the combustor 12 includes an ignition device (not shown), mixes the working fluid compressed by the high-pressure compressor 11H with the fuel, and burns the mixed gas. The generated combustion gas is discharged to the high-pressure turbine 13H.
  • the high pressure turbine 13H is provided behind the combustor 12.
  • the high-pressure turbine 13H includes a moving blade fixed to the high-pressure shaft 16b and a stationary blade fixed to the outer wall of the high-pressure turbine 13H.
  • the moving blades and stationary blades of the high-pressure turbine 13H are alternately installed along the shaft 2, and both are arranged in the circumferential direction CD.
  • the combustion gas passes through the moving blades and stationary blades of the high-pressure turbine 13H while expanding. In the process of this passage, the combustion gas rotates the moving blades of the high-pressure turbine 13H, and this rotational force is transmitted to the high-pressure compressor 11H via the high-pressure shaft 16b. As a result, the moving blades of the high-pressure compressor 11H rotate, and the working fluid is compressed.
  • the low pressure turbine 13L is provided behind the high pressure turbine 13H.
  • the low-pressure turbine 13L includes moving blades fixed to the low-pressure shaft 16a and stationary blades fixed to the outer wall of the low-pressure turbine 13L.
  • the moving blades and stationary blades of the low-pressure turbine 13L are alternately installed along the shaft 2, and both are arranged in the circumferential direction CD.
  • the combustion gas discharged from the high-pressure turbine 13H passes through the moving blades and stationary blades of the low-pressure turbine 13L while expanding. In the process of this passage, the combustion gas rotates the moving blades of the low-pressure turbine 13L, and this rotational force is transmitted to the low-pressure compressor 11L via the low-pressure shaft 16a. As a result, the moving blades of the low-pressure compressor 11L rotate to compress the working fluid.
  • the low pressure shaft 16a is located inside the high pressure shaft 16b in the radial direction.
  • the low-pressure shaft 16a and the high-pressure shaft 16b are arranged coaxially with the shaft 2 as the center, and both are rotatably supported by a support member such as a bearing (not shown).
  • the low pressure shaft 16a connects the low pressure compressor 11L (the moving blade of the low pressure compressor 11L) and the low pressure turbine 13L (the moving blade of the low pressure turbine 13L).
  • the high-pressure shaft 16b connects the high-pressure compressor 11H (the moving blade of the high-pressure compressor 11H) and the high-pressure turbine 13H (the moving blade of the high-pressure turbine 13H).
  • the core nozzle 14 is provided on the downstream side of the low pressure turbine 13L.
  • the core nozzle 14 is an annular flow path composed of a rearmost portion of the core case 15 and a tail cone 17 provided at the center thereof.
  • the core nozzle 14 discharges the combustion gas flowing out of the low-pressure turbine 13L toward the rear of the core engine 10.
  • the fan 20 has a moving blade (fan moving blade) 21 and a fan case 22.
  • the moving blades 21 are attached to the fan rotor 23 and are arranged radially around the shaft 2.
  • the fan rotor 23 is connected to the low pressure shaft 16a. As the low-pressure shaft 16a rotates, the rotor blade 21 and the fan rotor 23 rotate integrally. Due to the rotation of the rotor blade 21, the working fluid flows into the nacelle 24 from the outside of the engine 1, and a part of the working fluid is introduced into the core flow path 18 in the core case 15.
  • the fan case 22 is a hollow cylindrical member extending along the shaft 2 and surrounds a row of moving blades 21 (moving blade rows). That is, the maximum diameter of the fan case 22 is set to a value larger than the diameter of the circle including the chips of the plurality of rotor blades 21.
  • the length of the fan case 22 along the shaft 2 has at least a length for accommodating the moving blade 21, the portion on the upstream side of the core case 15, and the stationary blade 30. That is, the fan case 22 not only accommodates the moving blades (fan moving blades) 21, but also accommodates a part of the core engine 10 provided behind the moving blades 21, and bypasses the flow with the core case 15.
  • the road 25 is defined.
  • the fan case 22 is mounted and housed in the nacelle 24. Further, a stationary blade (fan stationary blade) 30 is provided in the bypass flow path 25.
  • the engine 1 (in other words, the fan 20) includes the stationary blade (fan stationary blade) 30 according to the present embodiment.
  • the stationary blades 30 are arranged in the circumferential direction CD to regulate the flow of the working fluid discharged from the moving blades 21.
  • the stationary blade 30 is located behind the moving blade 21 and extends from the outer surface 15b of the core case 15 to the inner surface 22a of the fan case 22.
  • the stationary blade 30 is provided in the bypass flow path 25 as, for example, an outlet guide blade (OGV).
  • OOV outlet guide blade
  • the hub 30a of the stationary blade 30 is attached to the outer surface 15b of the core case 15, and the tip 30b of the stationary blade 30 is attached to the inner surface 22a of the fan case 22.
  • the hub 30a and the tip 30b of the vane 30 may be supported by other corresponding structural members.
  • FIG. 2 is a diagram showing a stationary wing 30 according to the present embodiment.
  • FIG. 2 is a development view of the circumferential CD.
  • FIG. 3 is a diagram showing the relationship between the inclination of the moving blade 21 and the position of the maximum blade thickness of the stationary blade 30 (hereinafter, the maximum blade thickness position) M with respect to the axial direction AD, and FIG. 3A is shown in FIG.
  • FIG. 3B is a cross-sectional view taken along the line IIIB-IIIB (cross-sectional view on the hub side) shown in FIG. That is, FIG.
  • the stationary blade 30 has a blade body 32 having an airfoil cross section 31 shown in FIG.
  • the stationary blades 30 are arranged in the circumferential direction CD with a predetermined pitch P.
  • the wing body 32 has a leading edge 32a, a trailing edge 32b, and a dorsal (negative pressure surface) 32c and a ventral (positive pressure surface) 32d extending from the leading edge 32a to the trailing edge 32b.
  • the dorsal side 32c is a convex surface that is generally curved toward the rotation direction RD of the rotor blade 21 (see FIG. 3) (toward the front of the circumferential CD).
  • the ventral side 32d is also a concave surface substantially curved toward the rotation direction RD (see FIG. 3) of the moving blade 21 (toward the front of the circumferential direction CD). That is, the dorsal side 32c and the ventral side 32d are both curved in the same direction.
  • the airfoil cross section 31 of the blade body 32 satisfies the following conditions at least on the chip 30b side.
  • the blade body 32 has an airfoil cross section 31 that satisfies the following conditions at least on the chip 30b side.
  • the maximum blade thickness position M of the blade body 32 is set on the deployment surface of the circumferential CD in which the stationary blades 30 are arranged. (A) Located in the first region 33 near the trailing edge 32b of the wing body 32 from the intersection IP of the line 26 and the wing body 32, and (B) Located within the second region 34 with a code ratio of 0.2 to 0.8.
  • the maximum blade thickness position M is located in the third region 35 where the first region 33 and the second region 34 overlap.
  • the line 26 is a virtual line 26 passing through the leading edge 40a of another stationary blade 40 parallel to the extension line 28 of the camber line (airfoil center line) 27 of the moving blade 21 at the trailing edge 21b and adjacent to the circumferential CD. It is a line (see FIG. 3).
  • the extension line 28 is a tangent to the camber line 27 at the trailing edge 21b and extends posteriorly from the trailing edge.
  • the cord ratio is a value obtained by dividing the distance from the leading edge 32a of the blade body 32 to an arbitrary position on the cord of the blade body 32 by the cord length of the blade body 32.
  • the other stationary blade 40 described above is one of the plurality of stationary blades 30 arranged in the circumferential direction CD, and is located in front of the circumferential direction CD by the pitch P from the stationary blade 30 of interest. ..
  • the lower and upper limits of the code ratio in the second region 34 are such that the distance from the maximum blade thickness position M to the leading edge 32a or the distance from the maximum blade thickness position M to the trailing edge 32b is extremely short. It is set to suppress the induction of peeling caused by.
  • the maximum blade thickness position M of the present embodiment is determined by the above conditions. It is shifted to the trailing edge 32b side from the maximum blade thickness position of the conventional stationary blade. Further, if the maximum blade thicknesses of these airfoil cross sections are the same, the leading edge 32a of the present embodiment is sharper than the blunt leading edge formed on the conventional stationary blade. That is, according to the present embodiment, as compared with the conventional stationary blade, the thin portion 36 formed in the vicinity of the leading edge 32a expands from the leading edge 32a toward the trailing edge 32b.
  • a dual pole sound source or a quadrupole sound source can be assumed.
  • a double-pole sound source or a quadrupole sound source is a pressure vibration source generated by turbulence of a flow such as a wake or a vortex. It is known that the sound pressure level of the dual pole sound source and the sound pressure level of the quadrupole sound source are proportional to the sixth power and the eighth power of the flow velocity of the working fluid, respectively.
  • the working fluid flowing in the vicinity of the leading edge 32a of the stationary blade 30 can be decelerated and the sound pressure can be reduced by enlarging the portion 36 having a thin blade thickness based on the above two conditions.
  • FIG. 4 is a graph showing the distribution of wing surface Mach numbers on each dorsal side and each ventral side of the stationary blade 30 according to the present embodiment and the conventional stationary blade.
  • the vertical axis shows the wing surface Mach number
  • the horizontal axis shows the above-mentioned code ratio.
  • the solid line shows the blade surface Mach number distribution of the stationary blade 30 according to the present embodiment
  • the dotted line shows the blade surface Mach number distribution of the stationary blade of the conventional example.
  • the " ⁇ % span” in the figure refers to the distance from the hub (base) along the span direction with respect to the span length. Therefore, 90% span, 50% span, and 10% span mean the positions near the tip of the wing body, the center of the wing body, and the vicinity of the hub of the wing body, respectively.
  • FIG. 4A shows the blade surface Mach number distribution at 90% span on each dorsal side and each ventral side of the stationary blade 30 according to the present embodiment and the conventional stationary blade.
  • the number of blade surface Mach numbers in the vicinity of the leading edge 32a of the stationary blade 30 according to the present embodiment is smaller than that of the conventional stationary blade. This decrease appears on both the dorsal side 32c and the ventral side 32d, and according to computational fluid dynamics (CFD) analysis, it is shown that the sound pressure distribution near the leading edge 32a also decreases as compared with the conventional example. rice field. That is, according to the present embodiment, by defining the maximum blade thickness position that satisfies the above two conditions, noise can be reduced as compared with a stationary blade that does not satisfy the conditions.
  • CFD computational fluid dynamics
  • FIGS. 4 (b) and 4 (c) the number of wing surface Mach numbers in the vicinity of the leading edge 32a is also smaller than that of the conventional stationary blade at the center of the blade body 32 and on the hub side. It turns out.
  • Computational fluid dynamics (CFD) analysis also showed a decrease in sound pressure level, similar to the results for the 90% span. Therefore, the sound pressure level can be further reduced by forming the airfoil cross section 31 satisfying the above conditions over the entire span direction of the blade body 32.
  • FIG. 5 is a graph showing the numerical analysis results of each sound pressure level of the stationary blade 30 according to the present embodiment and the stationary blade of the conventional example.
  • the stationary blade 30 assumed in FIG.
  • FIG. 5 has an airfoil cross section 31 that satisfies the above conditions over the entire span direction.
  • FIG. 5 shows the sound pressure level of harmonics of the blade passing frequency (BPF), which is a component of the moving and stationary blade interference sound.
  • BPF blade passing frequency
  • the left side of the figure shows the comparison result of the sound pressure level in front of the stationary wing (in other words, the sound pressure level of the front sound), and the right side in the figure shows the sound pressure level behind the stationary wing (in other words, the rear).
  • the comparison result of the sound pressure level of the sound is shown. As shown in this figure, the sound pressure level is decreasing both in front of and behind the vane.
  • the flow velocity of the working fluid is larger on the tip 30b side of the moving blade 21 than on the hub 30a side of the moving blade 21. Therefore, by expanding the portion 36 on the tip 30b side of the stationary blade 30 toward the trailing edge 32b rather than the portion 36 on the hub 30a side, the separation of the working fluid on the hub 30a side is suppressed and the tip 30b side. It is possible to promote the suppression of noise.
  • the stationary blade according to the present embodiment can be applied to any one of a low-pressure compressor 11L, a high-pressure compressor 11H, a high-pressure turbine 13H, and a low-pressure turbine 13L. That is, at least one of these rotating machines may include a moving blade and a stationary blade provided behind the moving blade and having a blade body satisfying the conditions (a) and (b).

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Geometry (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

Une pale statique (30) disposée à l'arrière d'une pale de rotor (21) comprend un corps principal (32) de pale ayant une section transversale (31) en forme de pale. La position d'épaisseur maximale (M) de pale du corps principal (32) de pale dans la section transversale (31) en forme de pale remplit les conditions suivantes sur au moins le côté pointe (30b) du corps principal (32) de pale : (a) dans un plan développé dans la direction circonférentielle (CD) dans laquelle la pale statique (30) est disposée, la position est dans une première zone proche d'un bord de fuite (32b) du corps principal (32) de pale, à partir d'une intersection (IP) entre le corps principal (32) de pale et une ligne (26) parallèle à une ligne passant par une ligne d'extension (28) d'une ligne moyenne (7) de la pale de rotor (21) dans le bord arrière (21b) et passant par un bord d'attaque (40a) de l'autre pale statique (40) adjacente à la direction circonférentielle (CD) ; et (b) la position est dans une seconde zone (34) ayant un rapport de corde de 0,2 à 0,8.
PCT/JP2021/007122 2020-04-01 2021-02-25 Pale statique et turbine à gaz d'aéronef WO2021199802A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
JP2022511663A JP7294528B2 (ja) 2020-04-01 2021-02-25 静翼及び航空機用ガスタービンエンジン
EP21781375.7A EP4130436A4 (fr) 2020-04-01 2021-02-25 Pale statique et turbine à gaz d'aéronef
US17/808,114 US11867090B2 (en) 2020-04-01 2022-06-22 Stator vane and aircraft gas turbine engine

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US11867090B2 (en) 2024-01-09
JPWO2021199802A1 (fr) 2021-10-07
US20220316351A1 (en) 2022-10-06
EP4130436A4 (fr) 2024-04-24
JP7294528B2 (ja) 2023-06-20
EP4130436A1 (fr) 2023-02-08

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